In the case of hypersonic aircraft, the nose cones and leading edges of aerodynamic surfaces are subject to intense non-uniform atmospheric heating. The peak heating rate occurs at the stagnation point of flow with respect to the leading edge and decreases to a modest fraction of the peak flux at 90 degrees from the stagnation point, assuming the leading edge is a circular arc. Ablative cooling is the usual process utilized in present reentry vehicles, such as the Space Shuttle, to avoid the otherwise dangerous problem of overheating. The extreme heating during reentry is conventionally radiated to space by means of insulating tiles that are surface affixed with consequent limitations in the heating rates due to the maximum temperatures that can be withstood by the tiles. To compensate for the limited peak temperatures and, thus, limited heating rates of the insulating tiles, the Space Shuttle must reenter the earth's atmosphere at a large angle of attack.
For some missions envisioned for transatmospheric hypersonic vehicles, the radar signature associated with reentering at a large angle of attack may limit the mission effectiveness. Also, current mission scenarios envision multiple missions per day for such vehicles. Since the current history of insulating tile protection systems appears incompatible for multiple mission applications, there is a desire to investigate the possibility of utilizing active cooling to dissipate the heating on hypersonic vehicle leading edges and nose cones.
While the actual manner of successfully-utilizing active cooling has been unknown, it has been recognized that active cooling must take into account factors of critical importance. These include the capability of handling the circumferential variability of the heat flux; the fact that the angle of attack may change and cause the location of the highest heating rate to change on the circumference; the fact that there will be high intensity heating during acceleration to peak altitudes which, if vaporization of a working fluid is utilized, will have a tendency to blanket the leading edge surface with vapor since the acceleration field will be such that it will tend to urge the heavier liquid component away from the leading edge surface; and the fact that the peak intensities of heat flux are such that large temperature gradients across the wall comprising the leading edge may occur. In addition to factors of this type, it is generally recognized that the thermal stresses in the vehicle structure may limit the actual number of operating cycles that can satisfactorily be sustained by the vehicle.
Among the attempts to overcome the problems of aerodynamic heating is that disclosed in Billig et al U.S. Pat. No. 3,369,782. It is there proposed to provide an internally cooled leading edge structure for hypersonic vehicles in the form of a plurality of stainless steel tubes that are aligned in parallel relationship and bonded together with the corner portions of the tubes being aligned in clusters which are disposed in generally vertical planes perpendicular to the bonded portion of the tubes. With this arrangement, a coating of metal is electroformed over the tubes to form an aerodynamic wedge-shaped structure to be attached to the main body of a hypersonic airfoil.
Rice et al U.S. Pat. No. 2,941,759 discloses a heat exchanger construction of a very different type having porous walls and utilizes a coolant fluid which is uniformly distributed. The flow of the coolant fluid may be controlled and, if desired, completely dammed in selected areas. For this purpose, a cellular core which separates a porous skin and an inner sheet serves as a fluid conductor to provide fluid under pressure at defined, localized areas of the skin.
Still other heat exchange proposals include the rocket nozzle construction disclosed in Tumavicus U.S. Pat. No. 3,086,358, and the method of fabricating hollow structures having cooling channels disclosed in Dederra et al U.S. Pat. Nos. 3,690,103 and 3,692,637, as well as the methyl cooling system disclosed in McLauchlan U.S. Pat. No. 3,103,885.
Despite these attempts to provide a cooling unit, it has remained to successfully accomplish the objective of satisfactorily cooling a nose cone or a leading edge in a manner that avoids transmittal of thermal bending loads. Accordingly, the present invention is directed to overcoming the above-stated problems and accomplishing the stated objects by providing an active structural cooling unit for an aircraft leading edge or nose cone.